Integrated system for controlling a reciprocating internal combustion aircraft engine

Abstract

An integrated system for controlling a reciprocating internal combustion aircraft engine comprises an electronic control unit having two pairs of microprocessors, each pair forming a control channel. The system is designed to be capable of receiving signals from engine sensors, including vibration and acoustic emission sensors, a set of external engine parameter sensors and a set of internal engine operating parameter sensors. The system compares the values of the signals received from each of the sensors with a corresponding set of threshold value ranges, assesses the operating state and residual life of the engine on the basis of the values of the signals from said sensors, and generates a control signal for each of the actuating mechanisms of the engine, which is sent over the two control channels. The technical result consists in more accurate and reliable control of a reciprocating internal combustion aircraft engine.

Description

[0001] INTEGRATED CONTROL SYSTEM FOR AIRCRAFT PISTON INTERNAL COMBUSTION ENGINES

[0002] FIELD OF TECHNOLOGY TO WHICH THE INVENTION RELATES

[0003] The present technical solution relates to the field of internal combustion engine control, in particular to an integrated control system for an aircraft piston internal combustion engine.

[0004] LEVEL OF TECHNOLOGY

[0005] The prior art discloses information source US 2012 / 0192833 A1, published on August 2, 2012, which discloses a control device for an internal combustion engine. In the prior art device, an operating status value determination unit detects two or more operating status values ​​indicating the operating status of the internal combustion engine. A filter processing unit applies filter processing to the detected operating status values, and an operating status difference calculation unit calculates the difference between the filter-processed operating status value and the corresponding unfiltered operating status value to calculate two or more operating status value differences.The operation status value difference normalization unit normalizes two or more operation status value differences based on predetermined reference values ​​for the two or more operation status values ​​to calculate two or more normalized operation status value differences; and the transient correction unit corrects the control quantity for controlling the output of the internal combustion engine based on it when the engine is in the transient operation mode.

[0006] In addition, the prior art discloses information source US ! 2016 / 0223422 A1, published on 04.08.2016, which discloses a method for monitoring an operating event of an internal combustion engine, which includes receiving a noise signal sensed by a knock sensor, :located in or near an internal combustion engine, correlating a noise signal with a signature having at least an ADSR envelope indicative of an operational event, and detecting whether an operational event has occurred based on the correlation of the noise signal with the signature.

[0007] Finally, the prior art discloses information source US 2019 / 0128200 A1, published on 02.05.2019, which discloses a piston engine system that includes a cylinder, a piston located within the cylinder, a knock sensor located near the cylinder and configured to detect vibrations of the cylinder, the piston, or both, a crankshaft sensor configured to detect the angle of rotation of the crankshaft, and a controller communicatively connected to the knock sensor and the crankshaft sensor.The controller is configured to receive a raw detonation signal from the detonation sensor and a crankshaft rotation angle signal from the crankshaft sensor corresponding to vibrations of the cylinder, piston or both relative to the crankshaft rotation angle, converting the raw detonation signal into a digital signal of a value and at least one of the crankshaft rotation angle for the start of combustion, peak ignition pressure, percentage of combustion of the mass fraction of fuel or a combination thereof based on the digital signal of the value and the crankshaft rotation angle.

[0008] A common drawback of the known state-of-the-art solutions is the limited ability to control the operation of an internal combustion engine, which is based primarily on the use of vibration or detonation sensors, which can lead to erroneous or insufficiently accurate decisions when controlling other types of internal combustion engines.

[0009] Thus, the solutions known in the prior art do not allow for the implementation of complex integrated control of an aircraft piston internal combustion engine, which requires taking into account a multitude of both external and internal factors related to the operating process of an aircraft piston internal combustion engine.

[0010] The proposed solution differs from the solutions known from the prior art in that the control of an aircraft piston internal combustion engine is carried out in an integrated manner using two control channels, each with a pair of microprocessors that control the engine based on the current operating mode of the engine, the estimated state of performance and the predicted remaining life of the engine using a set of signals from vibration sensors, acoustic emission sensors, a set of sensors of external engine parameters and a set of sensors of internal engine operating parameters.

[0011] ESSENCE OF THE INVENTION The technical problem, which the claimed technical solution is aimed at solving, is the creation of a control system for an aircraft piston internal combustion engine, which will ensure efficient control of an aircraft piston internal combustion engine.

[0012] The technical result achieved by solving the above technical problem is an increase in the accuracy and reliability of control of an aircraft piston internal combustion engine due to the integrated analysis of signals from vibration and acoustic emission sensors together with signals from sensors of a different physical nature, data on the operating modes and state of the engine with duplication of control channels for the engine actuators.

[0013] According to one embodiment of the invention, an integrated control system for an aircraft piston internal combustion engine is proposed, comprising: an electronic control unit comprising two pairs of microprocessors, each of which forms a control channel and is configured to: receive signals from engine sensors, including vibrometry sensors, acoustic emission sensors and a set of sensors of external engine parameters and a set of sensors of internal engine operating parameters; compare the value of the received signals of each of the sensors with a corresponding set of threshold value ranges; evaluate the state of performance and predict the remaining life of the engine based on the values ​​of the signals of the vibrometry sensors, acoustic emission sensors and a set of values ​​of the received signals of the set of sensors of external engine parameters and the set of sensors of internal engine operating parameters;generating a control signal for each actuator of the engine based on the current operating mode of the engine, the results of the comparison, the estimated state of performance and the predicted remaining life of the engine; and sending the generated control signals to each actuator via two control channels.

[0014] In a particular embodiment of the proposed system, each range of threshold values ​​determines the operating mode of the engine.

[0015] In a particular embodiment of the proposed system, the engine actuators include a throttle valve, injection nozzles, and an ignition system. In a particular embodiment of the proposed system, the set of external engine parameter sensors comprises at least one of an ambient air pressure sensor, an ambient air temperature sensor, an ambient air humidity sensor, an aircraft center of gravity overload sensor, an aircraft angular velocity sensor, a propeller blade position sensor, and an exhaust temperature sensor.

[0016] In a particular embodiment of the proposed system, the set of sensors for internal engine operating parameters comprises at least one of an intake manifold pressure sensor, a crankshaft speed sensor, a throttle position sensor, and a fuel line pressure sensor.

[0017] DESCRIPTION OF DRAWINGS

[0018] The invention will be further described in accordance with the accompanying drawings, which are provided to illustrate the invention and in no way limit its scope. The following drawings are attached to the application:

[0019] Fig. 1 is a schematic block diagram of an exemplary integrated control system for an aircraft piston internal combustion engine according to one embodiment of the present invention;

[0020] Fig. 2 illustrates the arrangement of sensors for assessing the technical condition of an engine section according to one embodiment of the present invention;

[0021] Fig. 3 illustrates an example of analysis of vibration data of engine operation in different modes according to one embodiment of the present invention;

[0022] Fig. 4 illustrates the decision making regarding an unacceptable vibration level for one of the analysis frequencies according to one embodiment of the present invention.

[0023] DESCRIPTION OF THE INVENTION

[0024] The following detailed description of the invention includes numerous implementation details intended to provide a clear understanding of the present invention. However, one skilled in the art will readily appreciate how the present invention may be utilized with or without these implementation details. In other instances, well-known methods, procedures, and components have not been described in detail to avoid unnecessarily obscuring the features of the present invention.

[0025] Furthermore, it will be clear from the foregoing description that the invention is not limited to the embodiment described. Numerous possible modifications, changes, variations, and substitutions, while preserving the spirit and form of the present invention, will be apparent to those skilled in the art.

[0026] The proposed solution is an integrated control system for an aircraft piston internal combustion engine. In a specific embodiment, the system may be referred to as an integrated safety and airworthiness system (ISAS).

[0027] The integrated control system according to the present invention (hereinafter referred to as the control system or system) includes engine sensors and an electronic control unit. In a particular embodiment, the electronic control unit may be a full authority digital engine control (FADEC). In Fig. 1, which shows an exemplary integrated control system for an aircraft piston internal combustion engine, the electronic control unit is designated as an ISAS FADEC.

[0028] Engine sensors (aircraft piston internal combustion engine) include vibration sensors, acoustic emission sensors, a set of sensors for external engine parameters, and a set of sensors for internal engine operating parameters.

[0029] The set of external engine parameter sensors includes at least one of an ambient air pressure sensor, an ambient air temperature sensor, an ambient air humidity sensor, an aircraft center of gravity overload sensor, an aircraft angular velocity sensor, a propeller blade position sensor, and an exhaust temperature sensor. The set of external engine parameter sensors may be part of an aircraft structural condition monitoring system.

[0030] The set of internal engine performance sensors includes at least one of an intake manifold pressure sensor, a crankshaft speed sensor, a throttle position sensor, and a fuel line pressure sensor. Additionally, the set of internal engine performance sensors includes oil condition sensors.

[0031] Fig. 2 shows an approximate arrangement of sensors for assessing the technical condition of a cross-section of an engine according to one particular embodiment of the invention. Pos. 1 - arrangement of temperature sensors, Pos. 2 - arrangement of vibration sensors, Pos. 3 - arrangement of oil condition sensors, Pos. 4 - arrangement of acoustic emission sensors.

[0032] The electronic control unit is implemented on a board containing two pairs of microprocessors, each of which forms a control channel. In Fig. 1, the first pair of processors consists of MCU #1 (RTOS) and MCU #3 (HAL) and forms Control Channel 1, while the second pair of processors consists of MCU #2 (RTOS) and MCU #4 (HAL) and forms Control Channel 2.

[0033] The electronic control unit is housed in a protected housing.

[0034] Four microprocessors perform identical control functions but use different implementation systems. The first two microprocessors (#1 and #2) operate under the RTOS operating system, which enables time-based management of all system resources and redundant computation. The third and fourth microprocessors control computations and operate in the DIRECT system mode.

[0035] Engine sensor data is transmitted via UART, I2C, or SPI data interfaces, depending on the required data transfer rate, and is then fed into the ECU's data bus. Once in the ECU, the data is duplicated and transmitted to different processing channels (Channel 1 and Channel 2). It's also worth noting that most sensors are duplicated, meaning one sensor is physically connected to Channel 1, and the other to Channel 2.

[0036] The control channels of the electronic control unit are connected to the actuators of an aircraft piston internal combustion engine. These control channels transmit signals controlling the operation of the actuators. The engine actuators include the throttle valve, fuel injectors, and ignition system. The proposed system implements two independent engine control channels through the control of the engine actuators. Figure 1 shows an engine control diagram with duplicate channels.

[0037] Redundant control of engine actuators via two control channels, using the throttle valve as an example, is achieved as follows: a gear is mounted on the throttle shaft, meshing with two other gears attached to the shafts of two independent servo drives. If one of the servo drives fails, the working servo drive continues to rotate the throttle shaft, as well as the shaft of the faulty servo drive. This arrangement provides redundant control of the throttle valve, guaranteeing its operation even if one servo drive fails. Redundant control of other engine actuators occurs in a similar manner.

[0038] Additionally, the integrated control system's electronic control unit (ECU) board contains non-volatile EEPROM memory, which stores tabular data on engine operating modes. An SSD drive, also located on the ECU board, is used to accumulate the history of data received from sensors and the history of engine operation in various modes. Access to this drive for viewing the data history is via a USB interface via a connection to the ECU on the ground.

[0039] An approximate embodiment of the functioning of the invention

[0040] The integrated control system supports several control modes for the operation of an aircraft piston internal combustion engine.

[0041] 1) Normal operating mode (NORMAL)

[0042] 2) Alternative operating mode (ALTERNATE)

[0043] 3) Reduced operating mode (DEGRADED)

[0044] 4) direct operating mode (DIRECT)

[0045] Normal operating mode (NORMAL) is the primary engine control mode in all operating modes. The system analyzes current and historical parameters and selects optimal engine operating modes. Data is not written to memory. PC17RU2024 / 000391

[0046] Alternate mode (ALTERNATE) is the primary engine control mode in a situation where the main computer has encountered a runtime error and the system has restarted in alternate mode. Data is written to memory.

[0047] Reduced operating mode (DEGRADED) - this mode is an emergency mode for the control unit. In this mode, the technical condition analysis system is disabled, and the engine management system uses a minimum of sensors to control the engine. Data is not written to memory.

[0048] Direct mode (DIRECT) operates from a single computing unit and performs only engine control with a minimal set of functions. No data is written to memory.

[0049] In addition, the engine itself operates in the following modes: start, warm-up, taxiing, maximum takeoff, maximum continuous, cruise, descent - the parameters of all modes are optimally selected according to the planned altitude and flight speed of the aircraft.

[0050] The integrated control system, via an electronic control unit, receives signals from engine sensors, including vibration sensors, acoustic emission sensors, a set of sensors for external engine parameters, and a set of sensors for internal engine operating parameters.

[0051] Fig. 3 shows an example of the analysis of vibration data of engine operation in different modes.

[0052] Next, the electronic control unit compares the received signals from each sensor with a corresponding set of threshold ranges. Each signal type has its own threshold range. Each threshold range determines the engine's operating mode.

[0053] Then, the electronic control unit evaluates the state of performance and predicts the remaining life of the engine based on the values ​​of the signals from the vibration and acoustic emission sensors and the set of values ​​of the received signals from the set of sensors of external engine parameters and the set of sensors of internal engine operating parameters.

[0054] Figure 4 shows an example of deciding on an unacceptable vibration level for one of the analysis frequencies, along with vibration value ranges for assessing the engine's operating condition. The control system, via an electronic control unit, processes data from various sensors in parallel. For each type of data, the obtained analysis results are aggregated and used to make decisions about the reliability of the obtained data through comparative analysis, taking into account information from sensors of a different physical nature.

[0055] For example, if the temperature in one of the cylinders begins to rise, the system compares it with the temperature in other cylinders and also analyzes data from other physical sensors to determine the cause of the temperature increase.

[0056] Data from acoustic emission sensors is used to detect the onset of failure of key structural elements.

[0057] Data from vibration sensors is a key source of information on engine condition and performance. However, to obtain the most accurate and comprehensive picture of engine condition, vibration sensors must be integrated with data from other sensors. This comprehensive integration allows for the creation of a more accurate model of engine performance, taking into account all its parameters and characteristics.

[0058] The residual life of the engine is estimated by analyzing wave vibrometry and acoustic emission sensors, knowing the current and past operating modes.

[0059] The failure safety assessment is carried out during engine operation and signals from sensors of systems that are critical for engine operation:

[0060] 1) temperature sensor;

[0061] 2) vibration sensors;

[0062] 3) intake manifold pressure sensors;

[0063] 4) crankshaft speed sensor;

[0064] 5) throttle position sensor;

[0065] 6) fuel line pressure sensor;

[0066] 7) ambient air pressure sensor;

[0067] 8) ambient air temperature sensor;

[0068] 9) ambient air humidity sensor;

[0069] 10) overload sensor in the center of gravity of the aircraft;

[0070] 11) aircraft angular velocity sensor;

[0071] 12) propeller blade position sensors;

[0072] 13) Exhaust temperature sensor. Based on the current engine operating mode, comparison results, estimated engine performance, and predicted remaining engine life, the electronic control unit generates a control signal for each engine actuator and sends the generated control signals to each actuator via two control channels.

[0073] During normal operation of the engine and all its systems, the NORMAL control system operating mode is implemented.

[0074] If the onset of a degradation process in the engine is identified, which is mechanical in nature and does not affect the control system, then the control system carries out a system hazard assessment and makes changes to the engine operating modes.

[0075] If a shutdown of a number of sensors or control system errors is identified, the system switches to other control modes with auxiliary functions disabled and focusing only on maintaining engine performance.

[0076] The integrated control system's control modes for the operation of an aircraft piston internal combustion engine ensure a specified level of engine failure safety by calculating the dynamic parameters of the combustion process, analyzing acoustic vibration parameters from sensors, and restoring and verifying the correct operation of the engine.

[0077] In addition, the integrated control system additionally contains controls

[0078] Engine operating mode is determined by the throttle valve opening angle. The engine management system determines the optimal ignition timing and regulates fuel delivery, ensuring the engine operates at the specified operating mode.

[0079] In unmanned aerial vehicles (UAVs), signals from the UAV's automatic control system (ACS) serve as engine controls. In manned vehicles, the pilot uses the throttle lever (throttle stick).

[0080] During operation and failure of sensors, the system is reconfigured to simple control laws with the loss of some functions, but at any time the necessary set of functions remains to maintain a controlled operating mode of the engine and propeller. io PC17RU2024 / 000391

[0081] To switch between different operating modes, information from built-in sensors is used, which control the operation of all systems.

[0082] During normal operation, all processors are active, and the system makes decisions about engine operation based on quotas. When the engine enters abnormal operating modes, the principle of saving the engine's primary functions is implemented.

[0083] As can be seen from the preceding description, when using the claimed integrated system, the problem of controlling an internal combustion engine with a fully autonomous electro-digital controller (electronic control unit) of the engine is solved, in which not only the typical FADEC functions are carried out to maintain the operation of the engine and meet the requirements of the control system, such as:

[0084] 1) throttle control;

[0085] 2) control of injection nozzles;

[0086] 3) ignition control, but also solves the problem of assessing the resource and monitoring the engine’s failure safety.

[0087] This enables fail-safe control of an aircraft piston engine, taking into account redundancy and backup. This includes an assessment of the engine's service life and recommendations for maintenance inspection intervals.

[0088] A combination of acoustic emission and vibration parameters, as well as cylinder temperature, exhaust, and inlet manifold pressure data, provides a reliable tool for assessing engine life and operating parameters. Triple electronic redundancy of electrical systems and dual redundancy of all critical mechanical systems ensures a level of failure safety in accordance with NLG BAS-ST and BAS-VT requirements.

[0089] The proposed integrated system generates control signals for the controlled object and also monitors the object's condition and issues corrective signals. These are not two separate systems, but one. This means that the problem is solved using information from all sensors in an integrated manner, taking into account instantaneous values ​​and accumulated data history. The system also utilizes various types of engine condition data (control system tabular data, data from vibration sensors, and other physical measuring instruments).li Thus, due to the integrated analysis of signals from vibration and acoustic emission sensors together with signals from sensors of a different physical nature, data on operating modes and engine condition, the proposed integrated control system for an aircraft piston internal combustion engine ensures an increase in control accuracy, and due to the duplication of control channels for the engine actuators, an increase in the reliability of control of the aircraft piston internal combustion engine is ensured.

[0090] The microprocessors of the electronic control unit provide the data processing necessary to implement the claimed solution by executing machine-readable commands contained in the RAM, which configure the microprocessors to perform the basic computing operations necessary for the operation of the electronic control unit or the functionality of one or more of its components.

[0091] Memory is typically implemented as RAM, where the necessary software logic is loaded to provide the required functionality. When implementing the proposed solution, the memory capacity required for its implementation is allocated.

[0092] The device components are connected via a common data bus.

[0093] In these application materials, a preferred disclosure of the implementation of the claimed technical solution was presented, which should not be used as limiting other, particular embodiments of its implementation that do not go beyond the scope of the requested scope of legal protection and are obvious to specialists in the relevant field of technology.

Claims

PCI7RU2024 / 000391 CLAUSES OF THE INVENTION 1. An integrated control system for an aircraft piston internal combustion engine, comprising: an electronic control unit containing two pairs of microprocessors, each of which forms a control channel and is configured to: receive signals from engine sensors, including vibration sensors, acoustic emission sensors and a set of sensors of external engine parameters and a set of sensors of internal engine operating parameters; compare the value of the received signals from each of the sensors with a corresponding set of threshold value ranges; evaluate the state of performance and predict the remaining service life of the engine based on the values ​​of the signals from the vibration sensors, acoustic emission sensors and a set of values ​​of the received signals from the set of sensors of external engine parameters and the set of sensors of internal engine operating parameters;generating a control signal for each actuator of the engine based on the current operating mode of the engine, the results of the comparison, the estimated state of performance and the predicted remaining life of the engine; and sending the generated control signals to each actuator via two control channels.

2. The system according to claim 1, wherein each range of threshold values ​​determines the operating mode of the engine.

3. The system of claim 1, wherein the engine actuators include a throttle valve, injection nozzles and an ignition system.

4. The system according to claim 1, in which the set of external engine parameter sensors comprises at least one of an ambient air pressure sensor, an ambient air temperature sensor, an ambient air humidity sensor, an overload sensor at the center of gravity of the aircraft, an angular velocity sensor of the aircraft, a propeller blade position sensor, and an exhaust temperature sensor.

5. The system of claim 1, wherein the set of internal engine operating parameter sensors comprises at least one of an intake manifold pressure sensor, a crankshaft speed sensor, a throttle position sensor, and a fuel line pressure sensor.

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